With reference to FIG. 1, a conventional high bypass ratio turbofan engine generally indicated at 10 has a principal and rotational axis 11. This turbofan engine comprises a three-shaft gas turbine engine having high, intermediate and low pressure spools. Such engines are typically provided with a single rotating stage in both the high and intermediate turbines. The low pressure turbine however, usually comprises many stages.
The engine 10 comprises, in axial flow series, an air intake 12, a propulsive fan 13, an intermediate pressure compressor 14, a high pressure compressor 15, combustion equipment 16, a high-pressure turbine 17, an intermediate pressure turbine 18, a low-pressure turbine 19 and a core engine exhaust nozzle 20. A nacelle 21 generally surrounds the engine 10 and defines the intake 12, a bypass duct 22 and a bypass exhaust nozzle 23. The fan 13 is circumferentially surrounded by a fan casing 30, which is supported by an annular array of outlet guide vanes 27.
The gas turbine engine 10 works in a conventional manner so that air entering the intake 11 is accelerated by the fan 13 to produce two air flows: a first air flow into the intermediate pressure compressor 14 and a second air flow which passes through a bypass duct 22 to provide propulsive thrust. The intermediate pressure compressor 14 compresses the air flow directed into it before delivering that air to the high pressure compressor 15 where further compression takes place. The compressed air exhausted from the high pressure compressor 15 is directed into the combustion equipment 16 where it is mixed with fuel and the mixture combusted. The resultant hot combustion products then expand through, and thereby drive the high, intermediate and low pressure turbines 17, 18, 19 before being exhausted through the nozzle 20 to provide additional propulsive thrust. The high, intermediate and low pressure turbines 17, 18, 19 respectively drive the high and intermediate pressure compressors 15, 14 and the fan 13 by interconnecting shafts 24, 25, 26 respectively thereby making up high, intermediate and low pressure spools.
Control of the gas turbine engine requires analysis of various data consisting of measured engine parameters converted into electrical signals by sensors sited on the engine. These parameters include shaft speeds, temperatures, pressures and fuel flow rate.
The measured pressures include the inlet pressure at the fan face (P20), the pressure at the exit of the intermediate compressor (P25), the high pressure compressor delivery pressure (P30), the hot nozzle pressure (P50), and the by-pass pressure (P160). The location of each of these pressure variables within a typical three-shaft high bypass ratio turbofan engine is shown in FIG. 1.
Of these pressures, P30 being the pressure at the exit from the high pressure compressor is the highest pressure in the engine. Due to the hostile environment at the exit from the high pressure compressor, the associated pressure sensor is remotely located with a pipe connecting it to the exit region.
The connecting pipe is known to fill with moist air and in icing conditions this can freeze such that the pipe is blocked. This has the effect of ‘freezing’ the static pressure measurement because all variation in the measured pressure ceases until the pipe unfreezes.
A known technique for detecting the freezing of the P30 signal pipe relies on comparing a P30 value modelled from other engine parameters (such as, for example, fuel flow rate or high pressure compressor delivery temperature) with the actual measured P30 value. When a significant difference appears between the modelled and measured values then a freezing of the P30 signal pipe may be indicated.
On detection of the freezing of the P30 signal pipe, the control strategies of the engine change from using the measured P30 value to using the modelled P30 value.
In such circumstances, if there is little or no change in engine conditions (for example during cruise of the aircraft) then a P30 pipe freeze event may remain undetected until the pilot requires a change in engine thrust. The frozen P30 signal pipe may then cause the engine to produce an incorrect thrust level (i.e. higher or lower than requested) which would represent a safety risk during this critical event.
In effect, it is a dormant failure waiting for a demanded change in engine conditions to make it evident.
A disadvantage of the prior art system is that while it would detect the frozen P30 signal pipe and activate a different control regime for the engine using the modelled P30, there remains the possibility that a significantly incorrect thrust level may exist for a short period of time but at a potentially critical one.
A further disadvantage of the prior art system is that an error message, highlighting the frozen P30 signal pipe, is provided to the pilot. This adds to the pilot's workload during a critical time and may distract the pilot from potentially more important issues.